This disclosure relates to a cooling passage configuration for a gas turbine engine airfoil.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
Many blades and vanes include internal cooling passages that extend radially. The cooling passage may structurally weaken the blade. For some turbine blade applications, ribs are used in the cooling passage to strengthen the blade. In one example, these ribs transverse the entire span of the blade between the pressure and suction sides all the way from the root of the blade to its tip. Multiple layered cores result in half width feed cavities in the root of the part which greatly reduce the amount of air available to any given airfoil cavity.
In non-layered cores, one typical solution to increase the amount of flow to a given airfoil cavity without eliminating the ribs entirely from the blade is to place a “riblet” in the root of the core only. This “riblet” is a rib that stops within the root or neck of the blade and does not continue to the airfoil. The height of these ribs is dictated by the stresses in the root. Adjacent cooling passageways remain discrete and fluidly separated from one another.